This invention generally relates to a gas turbine engine, and more particularly to a nacelle inlet for a turbofan gas turbine engine.
In an aircraft gas turbine engine, such as a turbofan engine, a fan delivers air to a compressor. The pressurized air is mixed with fuel in a combustor for generating hot combustion gases. The hot combustion gases flow downstream through a turbine stage, which extracts energy from the gas. The turbine powers the fan and compressor.
Combustion gases are discharged from the turbofan engine through a core exhaust nozzle and fan air is discharged through an annular fan exhaust nozzle defined at least partially by a nacelle surrounding the core engine. A majority of propulsion thrust is provided by the pressurized fan air which is discharged through the fan exhaust nozzle, while the remaining thrust is provided from the combustion gases discharged through the core exhaust nozzle.
It is known in the field of aircraft gas turbine engines that the performance on the turbofan engine varies during diverse flight conditions experienced by the aircraft. An inlet lip section located at the foremost end of the turbofan nacelle is typically designed to enable operation of the turbofan engine and prevent the separation of airflow from the inlet lip section of the nacelle during diverse flight conditions. For example, the inlet lip section requires a “thick” inlet lip section design to support operation of the turbofan during specific flight conditions, such as cross-wind conditions, take-off and the like. Disadvantageously, the “thick” inlet lip section may reduce the efficiency of the turbofan engine during other conditions, such as cruise conditions of the aircraft.
Accordingly, it is desirable to optimize the performance of a turbofan gas turbine engine during diverse flight requirements to provide a nacelle having a reduced thickness, reduced weight and reduced drag.